CN101583487A - Hybrid composite-metal aircraft landing gear and engine support beams - Google Patents
Hybrid composite-metal aircraft landing gear and engine support beams Download PDFInfo
- Publication number
- CN101583487A CN101583487A CNA2007800417978A CN200780041797A CN101583487A CN 101583487 A CN101583487 A CN 101583487A CN A2007800417978 A CNA2007800417978 A CN A2007800417978A CN 200780041797 A CN200780041797 A CN 200780041797A CN 101583487 A CN101583487 A CN 101583487A
- Authority
- CN
- China
- Prior art keywords
- metalwork
- composite
- interior
- metal
- tapered end
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64C—AEROPLANES; HELICOPTERS
- B64C25/00—Alighting gear
- B64C25/02—Undercarriages
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B15/00—Layered products comprising a layer of metal
- B32B15/04—Layered products comprising a layer of metal comprising metal as the main or only constituent of a layer, which is next to another layer of the same or of a different material
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B3/00—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form
- B32B3/02—Layered products comprising a layer with external or internal discontinuities or unevennesses, or a layer of non-planar shape; Layered products comprising a layer having particular features of form characterised by features of form at particular places, e.g. in edge regions
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/40—Arrangements for mounting power plants in aircraft
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64D—EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT
- B64D27/00—Arrangement or mounting of power plants in aircraft; Aircraft characterised by the type or position of power plants
- B64D27/40—Arrangements for mounting power plants in aircraft
- B64D27/402—Arrangements for mounting power plants in aircraft comprising box like supporting frames, e.g. pylons or arrangements for embracing the power plant
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B32—LAYERED PRODUCTS
- B32B—LAYERED PRODUCTS, i.e. PRODUCTS BUILT-UP OF STRATA OF FLAT OR NON-FLAT, e.g. CELLULAR OR HONEYCOMB, FORM
- B32B2605/00—Vehicles
- B32B2605/18—Aircraft
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/40—Weight reduction
Landscapes
- Engineering & Computer Science (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mechanical Engineering (AREA)
- Laminated Bodies (AREA)
- Manufacture Of Alloys Or Alloy Compounds (AREA)
Abstract
提供了一种混合式复合金属部件。该部件包括细长内金属件、绕所述内金属件的至少一部分设置的外金属件以及在所述内金属件与所述外金属件之间设置的复合材料。所述部件还可包括连接所述内金属件与外金属件的密封件与至少一个紧固件中的至少一个。所述内金属件与所述外金属件都可包括至少一个渐缩端或梯形端。所述内金属件与所述外金属件的所述渐缩端均可包括双梯形度。
A hybrid composite metal part is provided. The component includes an elongated inner metal piece, an outer metal piece disposed about at least a portion of the inner metal piece, and a composite material disposed between the inner metal piece and the outer metal piece. The component may also include at least one of a seal and at least one fastener connecting the inner and outer metal members. Both the inner metal piece and the outer metal piece may include at least one tapered or trapezoidal end. The tapered ends of the inner metal piece and the outer metal piece may each include a double taper.
Description
Technical field
[0001] embodiment of the present disclosure relates to the formation of hybrid composite-metal component, and more specifically, relates to the apparatus and method for that is used to form hybrid composite-metal aircraft landing gear and engine beam.
Background technology
[0002] in many application, especially in aviation, navigation, space flight and building industry, importantly provides to have such as the particular characteristic of intensity and have minimum or the parts of the quality of decrease at least.Undercarriage and engine beam are generally heavy metal structures.For example, the aircraft 100 with undercarriage 200 shown in Figure 1.Undercarriage 200 is generally speaking below cockpit area 150.The main landing gear 200 of Fig. 1 is arranged near aircraft wing 101 places.In Fig. 2, aircraft engine 102 is supported by the engine beam 201 near aircraft wing 101.The undercarriage that is made of metal provides the protection of necessity so that prevent by the collision that chip caused on runway.Simultaneously, use the advantage of metal to be to support or to retrain main load.Certainly, use the big shortcoming of metal to be to be used to the quality that realizes that these structure purposes are required.Because designing requirement needs light structures, therefore, typically, the method for producing that the undercarriage that is formed by metal and engine beam need is difficulty and other design problem in addition quite.
[0003] requirement of incompressible in undercarriage, crooked, torsional load and runway chip has produced the demand for new undercarriage design.But new undercarriage design must be satisfied standard-required have more lightweight.All utilize metal or all utilize the existing and emerging technology of composite construction that the limited capability of realizing these requirements is provided.That is, be lighter than metal structure on the composite construction weight but need be used for the mould or the instrument of costliness of its manufacturing and autoclave or the forcing press that is used for its curing process.In addition, composite construction is subject to impact injury and the impossible weight that supports whole aircraft.Therefore, although have shortcoming on the weight, metal is still the material that is used for undercarriage and selects.Therefore, the net weight of undercarriage remains a problem of aircraft industry.
[0004] for the requirement of engine beam to similar to the undercarriage designing requirement.Engine beam must provide enough supports so that effectively resist the various load that comprise pitching load and edge loading that caused by engine.As the situation of undercarriage, expectation reduces the weight of engine support structure as far as possible and reduces the structural capacity that realizes that its load requires not seriously.Therefore, existence is for the demand of the new engine beam that reduces quality.Prior art and emerging technology provide the limited capability of finishing described demand.Typically, engine supports is made of metal.The mould or the instrument of the costliness that metallic supports need not used in compound support member is made.Therefore, metal remains the material selection of engine beam.Thereby the weight of engine beam remains designer's a problem.
[0005] therefore, the equipment and the method that are provided for forming hybrid parts will be favourable, wherein said hybrid parts certain at least intensity of providing by the conventional metals parts has been provided and the constant weight at least that provides by composite material component on advantage.In addition, the equipment and the method that are provided for forming as lower member will be favourable, and wherein said parts have reduced the gross weight of aircraft or other delivery vehicles and do not damaged its structural integrity.Be accompanied by littler construction weight, aircraft and other vehicles can carry bigger payload and realize bigger fuel efficiency.
Summary of the invention
[0006] embodiment of the present disclosure can satisfy above demand and obtain other advantage by the apparatus and method for that is provided for forming such as the hybrid composite-metal component of hybrid composite-metal aircraft landing gear and engine beam.In general, embodiment of the present disclosure is provided for forming the apparatus and method for of hybrid composite-metal component and need not cutter or autoclave processing, also can benefit from the performance and the characteristic of composite and metal material simultaneously.Particularly, hybrid composite-metal component can be formed by the metalwork that the composite that solidifies connects together, and wherein the composite of this curing occupies the gap between these metalworks.
[0007] in one embodiment, hybrid composite-metal component comprise elongated in metalwork (elongate inner metal piece), at least a portion setting of metalwork in this outer metalwork and in this metalwork with this outside the composite that is provided with between the metalwork.Be somebody's turn to do interior metalwork and should all can have relative convergent and non-tapered end by outer metalwork.Can be by the length that distance limited with approximately identical to the length that distance limited of non-tapered end by the tapered end of outer metalwork from the tapered end of interior metalwork to non-tapered end.Interior metalwork and outer metalwork can by in seal and at least one securing member at least one and be connected, wherein said securing member can be interior metalwork and outside the bolt that extends between the metalwork or around one section evenly spaced a plurality of securing member of outer metalwork.Interior metalwork can be formed by titanium with outer metalwork.Composite can be by forming with the graphite of resin dipping.Interior metalwork can comprise double taper or double tapered with the tapered end of outer metalwork.Simultaneously, the tapered end of interior metalwork and outer metalwork can be aligned but not tapered end is aligned too.
[0008] in another embodiment, provide the method that forms hybrid composite-metal component.This method comprises: make gappedly in the metalwork therebetween thereby interior metalwork mated outside, fill at least a portion in this gap with composite, with is connected in metalwork and outer metalwork.Interior metalwork with can comprise utilization seal and at least one securing member of attaching being connected of outer metalwork at least one measure.At least one securing member of attaching can comprise at least one bolting to interior metalwork and outer metalwork, and fix a plurality of bolts around the even compartment of terrain of outer metalwork.At least a portion of filling the gap with composite is included in composite that deposition is done in the gap and the composite of doing with the resin dipping.This method also comprises cured composite material.The curing of composite can comprise to this composite and applies heat or radiation.Simultaneously, this method is exerted pressure to composite during can being included in composite material solidification.
[0009] in another embodiment, provide aircraft component.The outer metal tube that this aircraft component comprises interior metal tube, is provided with around at least a portion of interior metal tube, and interior metal tube and outside the composite that is provided with between the metal tube.As foregoing, interior metal tube all can have at least one tapered end with outer metal tube.Interior metal tube all can comprise double taper or double tapered with the tapered end of outer metal tube.
Description of drawings
[0010] so roughly described embodiment of the present disclosure, will make with reference to explanation at accompanying drawing now, wherein said accompanying drawing and not drawn on scale, and wherein:
[0011] Fig. 1 is the diagram of aircraft, this illustrate below passenger cabin area undercarriage and near the main landing gear of wing.
[0012] Fig. 2 is the diagram near the engine beam of aircraft engine and wing.
[0013] Fig. 3 is the perspective view of elongated interior metalwork.
[0014] Fig. 4 is the cutaway view of elongated interior metalwork, and metalwork has the outer metalwork that is provided with around the part of interior metalwork in this.
[0015] Fig. 5 is the cutaway view according to metalwork in embodiment elongated, in this metalwork have around the outer metalwork of the part setting of interior metalwork and interior metalwork and outside the composite that is provided with between the metalwork.
[0016] Fig. 6 is the cutaway view of the interior piston that is provided with of a part of metalwork in being illustrated in.
The specific embodiment
[0017] describes embodiment now below with reference to the accompanying drawings more comprehensively, in the accompanying drawings, show some but not all embodiment.In fact, these embodiment can multi-formly be embodied and should be interpreted as restriction to the embodiment disclosed herein with many; And, provide these embodiment so that the disclosure will satisfy applicable legal requiremnt.Same numbers is represented similar elements in institute's drawings attached.
[0018] provides and to be used to various application and for example to can be used as the undercarriage principal post of aircraft and the hybrid composite-metal component of bogie (truck) or engine beam.Hybrid composite-metal component comprises elongated interior metalwork (elongate inner metal piece) 10, and this elongated interior metalwork 10 can have tapered end 11 and relative non-tapered end 12 as shown in Figure 3.Metalwork 10 can be formed by various metals in elongated, and described metal for example comprises titanium.Metalwork 10 can be solid or hollow in elongated.It can be cylindrical as shown in Figure 6 but also can be other shape.Hybrid composite-metal component also comprises outer metalwork 20.In this regard, Fig. 4 illustrates the outer metalwork 20 with tapered end 21 and non-tapered end 22.Outer metalwork 20 generally is hollow and is cylindrical, this cylindrical interior diameter that has greater than the overall diameter of interior metalwork 10.Like this, outer metalwork 20 can be arranged to whole then around its part if not around interior metalwork 10.Outer metalwork 20 can be rendered as and not be columniform shape.Typically, thus the length of outer metalwork 20 can fit in the outer metalwork 20 more than or equal to the interior metalwork 10 of the length of interior metalwork 10.Outer metalwork 20 can be formed by various metals, and described metal for example comprises titanium.In this regard, interior metalwork 10 can be formed by identical or different metal with outer metalwork 20.The interior diameter of outer metalwork 20 usually greater than the overall diameter of interior metalwork 10 to limit gap 13 betwixt.
[0019] just as shown in FIG. 5, the gap 13 between metalwork 20 and the interior metalwork 10 is filled by composite 30 outside.Composite 30 can comprise various composites, for example the graphite of impregnating resin.Typically, fill gaps 13 with composite 30 and comprise composite fibre or other dried composites are installed in the gap 13, for example by coiling, braiding or manual the placement, and then with resin transfer in gap 13.In case composite 30 has been placed in the gap 13 and resin has been transferred in the gap 13, then can for example by radiation, come cured composite material 30 by heating.The piston 18 in Fig. 5 also illustrates and partly is arranged in the metalwork 10 and the part of piston 18 are set in the cylinder 19.Piston 18 can be used to assist resin transfer, and tension force (tension) for example is provided.Although illustrating, Fig. 5 only have a piston 18 partly to be arranged in the interior metalwork 10, but other embodiment can comprise two or more pistons 18 that are at least partially disposed in the interior metalwork 10, for example, by two pistons 18 in metalwork 10 opposite ends in partly being arranged on.
[0020] typically, composite 30 basic or complete filling gaps 13.The width in gap 13 depends on uses especially load requirement and difference.For example, thus bigger and heavier aircraft needs bigger thickness of composite material to provide necessary intensity to be applied to carry-on load by hard landing with opposing under the maximum gross situation.The surface of metal parts contact composite resin material can be corroded and be engaged by viscose glue and fill in so that high bond strength to be provided.Outer metalwork 20 also typically is connected by the securing member such as bolt 5 with interior metalwork 10.In one embodiment, for example outer metalwork 20 can be connected by a plurality of bolts 5 that circumferentially distribute around outer metalwork 20 surfaces with interior metalwork 10.Typically, around outer metalwork 20 circumferentially, bolt 5 is spaced apart in uniform mode, but if desired, bolt 5 also can be spaced apart brokenly.Can use the major diameter securing member, particularly reverse and edge loading in order to resist.Additionally or alternatively, outer metalwork 20 and interior metalwork 10 can sealed be connected.The sealing part is high temperature resistant seal typically, for example polyimides.The outer surface of the inner surface of outer metalwork 20 and interior metalwork 10 can have one deck polytetrafluoroethylene (PTFE) that is used to protect described two surfaces
After solidifying, can remove polytetrafluoroethylene (PTFE)
Additionally or alternatively, outer metalwork 20 can comprise threaded metal parts with interior metalwork 10.
[0021] in Fig. 6, outer metalwork 20 has double taper or double tapered 15.This double taper 15 is shown as two different taper angle T1, T2 that cross over conical section 21 in Fig. 6.As shown in, the tapering that the least significant end tapering limits taper angle T2 in other words is bigger than another tapering usually, promptly for the longitudinal axis that is limited by interior metalwork 10 or outer metalwork 20, the tapering of least significant end tapering or qualification taper angle T2 is in the angle bigger than another tapering.Double taper or double tapered 15 can be the loading environment that composite 30 provides expectation.
[0022] those skilled in the art will expect satisfying multiple remodeling and other embodiment of following condition, i.e. advantage with the instruction content that is presented in front specification and the accompanying drawing of described multiple remodeling under these embodiment and other embodiment.For example, the one or both in interior metalwork 10 and the outer metalwork 20 all need not to have tapered end 11 and also can have cylindrical or even flared end outwardly.And, although illustrated and described cylindrical interior metalwork 10 and cylindrical outer metalwork 20, but if in metalwork 10 can fit at least in part in the outer metalwork 20, then in metalwork 10 can have other shape of cross sections and interior metalwork 10 and outer metalwork 20 with one or both in the outer metalwork 20 and can have different shape of cross sections.Therefore, should be understood that the disclosure is not limited to disclosed specific embodiment, and will retrofit intentionally and other embodiment is also contained within the scope of claims protections.Though used concrete term at this, they only are used to general and descriptive sense and not for the purpose of restriction.
Claims (20)
1. hybrid composite-metal component, it comprises:
Metalwork in elongated;
The outer metalwork that at least a portion of metalwork is provided with in described; And
Be placed in the composite between described interior metalwork and the described outer metalwork.
2. hybrid composite-metal component according to claim 1, wherein said interior metalwork all has relative tapered end and non-tapered end with described outer metalwork.
3. hybrid composite-metal component according to claim 2, metalwork and outer metalwork are connected in wherein said, make described in the described tapered end of metalwork aim at the described tapered end of described outer metalwork and the described non-tapered end of described interior metalwork is aimed at the described non-tapered end of described outer metalwork.
4. hybrid composite-metal component according to claim 2, the described tapered end of wherein said interior metalwork and described outer metalwork includes double taper.
5. hybrid composite-metal component according to claim 1, wherein said composite comprises the graphite of impregnating resin.
6. hybrid composite-metal component according to claim 1, further comprise connect described in seal and at least one securing member of metalwork and outer metalwork one of at least.
7. hybrid composite-metal component according to claim 6, wherein said at least one securing member are included in the bolt that extends between described interior metalwork and the outer metalwork.
8. hybrid composite-metal component according to claim 6, wherein said at least one securing member comprise one section evenly spaced a plurality of securing member of metalwork outside described.
9. hybrid composite-metal component according to claim 1, wherein said interior metalwork comprises titanium spare.
10. hybrid composite-metal component according to claim 1, wherein said outer metalwork comprises titanium spare.
11. a method that forms hybrid metal parts, it comprises:
Thereby interior metalwork mated outside make to have the gap therebetween in the metalwork;
Fill at least a portion in described gap with composite;
Connect described interior metalwork and described outer metal; And
Solidify described composite;
Wherein at least a portion of filling described gap with composite is included in the composite that deposition is done in the described gap, and wherein said method also comprises subsequently and floods described dried composite with resin.
12. the method for formation mixed composite parts according to claim 11 wherein connects described interior metalwork and described outer metalwork and comprises at least one the measure of using in seal and at least one securing member of attaching.
13. the method for formation mixed composite parts according to claim 12, wherein at least one securing member of attaching comprises at least one bolting to described interior metalwork and described outer metalwork.
14. the method for formation mixed composite parts according to claim 12, wherein at least one securing member of attaching comprises the fixing a plurality of bolts in the even compartment of terrain of metalwork outside described.
15. the method for formation mixed composite parts according to claim 11 is wherein solidified described composite and is comprised to described composite and apply radiation.
16. the method for formation mixed composite parts according to claim 11 is wherein solidified described composite and is comprised to described composite and apply heat.
17. the method for formation mixed composite parts according to claim 11 further comprises and moves a piston that is configured in the described interior metalwork.
18. an aircraft component, it comprises:
Interior metal tube;
The outer metal tube that at least a portion of metal tube is provided with in described; And
Be placed in the composite between described interior metal tube and the described outer metal tube.
19. aircraft component according to claim 18, wherein said interior metal tube and described outer metal tube all have at least one tapered end.
20. aircraft component according to claim 19, the described tapered end of wherein said interior metal tube and described outer metal tube includes double taper.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US11/608,534 | 2006-12-08 | ||
| US11/608,534 US20080283667A1 (en) | 2006-12-08 | 2006-12-08 | Hybrid composite-metal aircraft landing gear and engine support beams |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN101583487A true CN101583487A (en) | 2009-11-18 |
Family
ID=39789166
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CNA2007800417978A Pending CN101583487A (en) | 2006-12-08 | 2007-12-07 | Hybrid composite-metal aircraft landing gear and engine support beams |
Country Status (6)
| Country | Link |
|---|---|
| US (1) | US20080283667A1 (en) |
| EP (1) | EP2097255A2 (en) |
| JP (1) | JP2010512268A (en) |
| CN (1) | CN101583487A (en) |
| CA (1) | CA2666595A1 (en) |
| WO (1) | WO2008118229A2 (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104875878A (en) * | 2014-02-27 | 2015-09-02 | 空中客车西班牙运营有限责任公司 | Aircraft with a nacelle-housed main landing gear |
Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB2473007B (en) | 2009-08-26 | 2012-11-21 | Messier Dowty Ltd | Apparatus comprising an end fitting connected to a body |
| JP5514527B2 (en) * | 2009-12-11 | 2014-06-04 | ナブテスコ株式会社 | Aircraft reaction link |
| WO2020008372A1 (en) | 2018-07-03 | 2020-01-09 | University Of Notre Dame Du Lac | Polymer/exfoliated nano-composite films with superior mechanical properties |
| EP3611390B1 (en) | 2018-08-16 | 2020-12-23 | Crompton Technology Group Limited | Composite structure having a tapered joint and a method for making the same |
| EP3792173B1 (en) | 2019-09-16 | 2022-04-27 | SKF Aerospace France | Fail-safe system intended for use in an aircraft |
Family Cites Families (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US2584832A (en) * | 1945-04-07 | 1952-02-05 | Bendix Aviat Corp | Method of making a trunnion for aircraft landing gear |
| US3416975A (en) * | 1965-05-04 | 1968-12-17 | Nasa | Etching of aluminum for bonding |
| US3623203A (en) * | 1970-03-24 | 1971-11-30 | Avco Corp | Reinforced structural members and method of making same |
| US4185472A (en) * | 1978-03-27 | 1980-01-29 | Celanese Corporation | Fiber reinforced composite shaft with metallic connector sleeves mounted by radial pin interlock |
| CA1107769A (en) * | 1978-05-10 | 1981-08-25 | Henry E. Wilson | Composite fibrous tube energy absorber |
| US4300439A (en) * | 1979-09-10 | 1981-11-17 | United Technologies Corporation | Ballistic tolerant hydraulic control actuator and method of fabricating same |
| DE3017336A1 (en) * | 1980-05-06 | 1981-11-12 | Paul Dr. 1000 Berlin Mader | Tubular body subjected to bending stress - has internal adhesive length of reinforcing fibres, forming strands adapted to body interior |
| DE3103646C2 (en) * | 1981-02-04 | 1984-03-29 | Aluminium-Walzwerke Singen Gmbh, 7700 Singen | Pressure vessels for storing and transporting gaseous fluids |
| US4581077A (en) * | 1984-04-27 | 1986-04-08 | Nippon Mining Co., Ltd. | Method of manufacturing rolled titanium alloy sheets |
| US4923751A (en) * | 1986-10-21 | 1990-05-08 | The United States Of America As Represented By The Administrator Of The National Aeronautics And Space Administration | Seamless metal-clad fiber-reinforced organic matrix composite structures, and process for their manufacture |
| FR2675563B1 (en) * | 1991-04-22 | 1993-08-27 | Aerospatiale | METHOD OF MECHANICAL ASSEMBLY OF A TUBE OF COMPOSITE MATERIAL AND A METAL PART AND ASSEMBLY THUS CARRIED OUT. |
| US5332049A (en) * | 1992-09-29 | 1994-07-26 | Brunswick Corporation | Composite drill pipe |
| JPH0960676A (en) * | 1995-08-25 | 1997-03-04 | Yokohama Rubber Co Ltd:The | Shock relaxing material and manufacture thereof |
| US6893733B2 (en) * | 2000-07-07 | 2005-05-17 | Delphi Technologies, Inc. | Modified contoured crushable structural members and methods for making the same |
| US6830223B1 (en) * | 2000-11-16 | 2004-12-14 | Tyee Aircraft | Force sensor rod |
| US6582172B2 (en) * | 2001-08-29 | 2003-06-24 | The United States Of America As Represented By The Secretary Of The Navy | Isolated mechanical fastening system |
| DE102004008523B4 (en) * | 2004-02-20 | 2007-02-01 | Liebherr-Aerospace Lindenberg Gmbh | Method for producing a printing cylinder and piston rod for actuators or shock absorbers and method for their production |
-
2006
- 2006-12-08 US US11/608,534 patent/US20080283667A1/en not_active Abandoned
-
2007
- 2007-12-07 CN CNA2007800417978A patent/CN101583487A/en active Pending
- 2007-12-07 EP EP07873665A patent/EP2097255A2/en not_active Withdrawn
- 2007-12-07 CA CA002666595A patent/CA2666595A1/en not_active Abandoned
- 2007-12-07 WO PCT/US2007/086744 patent/WO2008118229A2/en active Application Filing
- 2007-12-07 JP JP2009540494A patent/JP2010512268A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN104875878A (en) * | 2014-02-27 | 2015-09-02 | 空中客车西班牙运营有限责任公司 | Aircraft with a nacelle-housed main landing gear |
| CN104875878B (en) * | 2014-02-27 | 2019-08-13 | 空中客车西班牙运营有限责任公司 | Aircraft with the main landing gear being placed in nacelle |
Also Published As
| Publication number | Publication date |
|---|---|
| CA2666595A1 (en) | 2008-10-02 |
| EP2097255A2 (en) | 2009-09-09 |
| US20080283667A1 (en) | 2008-11-20 |
| JP2010512268A (en) | 2010-04-22 |
| WO2008118229A2 (en) | 2008-10-02 |
| WO2008118229A3 (en) | 2009-03-19 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| US8205825B2 (en) | Engine pylon made from composite material | |
| RU2602100C2 (en) | Aircraft structure to ensure high resistance to composite side girder pulling out | |
| CN101583487A (en) | Hybrid composite-metal aircraft landing gear and engine support beams | |
| EP3076048B1 (en) | Composite actuator rod end assembly | |
| US9950779B2 (en) | Elongated structures and related assemblies | |
| US10501163B2 (en) | Pressure bulkhead for an aircraft fuselage, and an aircraft comprising such a pressure bulkhead | |
| US20100006698A1 (en) | Hybrid strut having metal and composite portions | |
| CA1084028A (en) | Main connector for an airfoil or wing | |
| FR2921899A1 (en) | METHOD FOR LOCALLY REINFORCING COMPOSITE MATERIAL ELEMENT AND CENTRAL BOAT BOILER FOR REINFORCED AIRCRAFT | |
| US8418962B2 (en) | Distribution of point loads in honeycomb panels | |
| KR20170113329A (en) | A helicopter with a fuselage and a composite tail boom | |
| CA2659448C (en) | Wing panel structure | |
| WO1998015455A1 (en) | Prestressed structure for aircraft and the method of producing the same | |
| US10272991B2 (en) | Metallic composite joint | |
| US10562650B2 (en) | Corrugated payload adaptor structure | |
| US20220170499A1 (en) | Structural fuse | |
| CN103350751B (en) | Jackstay energy-absorption device at lower part of passenger cabin floor of civil aircraft | |
| US9987772B1 (en) | System and method to manufacture composite structures | |
| Beam | Check for Preliminary Design Approach for Lattice Composite Keel Beam Fedir Gagauz and Serhii KryvendaⓇ National Aerospace University “Kharkiv Aviation Institute”, 17 Chkalova Street | |
| JP2022145620A (en) | Support structure for aircraft landing gear | |
| WO2023016761A1 (en) | Aerostructure component |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| C06 | Publication | ||
| PB01 | Publication | ||
| C10 | Entry into substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| C02 | Deemed withdrawal of patent application after publication (patent law 2001) | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20091118 |